1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, compressed air from a compressor is burned with a fuel in a combustor to produce a hot gas flow that is then passed through a turbine to produce mechanical energy by rotating the rotor shaft. In an aero engine, the rotor shaft drives the compressor and a bypass fan to power the aircraft. In an industrial gas turbine (IGT) engine, the rotor shaft drives an electric generator to produce electrical energy.
The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the highest temperature for which the turbine can be operated is dependent upon the material characteristics of the turbine parts, especially the first stage rotor blades and stator vanes or guides. These parts are exposed to the highest temperature flow from the combustor.
To allow for higher temperatures beyond the material properties of the turbine blades and vanes, these airfoils make use of complex internal cooling circuitry that provides a combination of convection cooling as well as impingement and film cooling of the inner airfoil surfaces and the outer airfoil surface. Modern airfoil cooling circuitry can allow for the operation of an airfoil under a temperature that exceeds the material melting temperature.
Cooling air for use in the airfoils is compressed air bled off from the compressor, and therefore the work used in compressing the cooling air for the airfoils is lost energy. Thus, the efficiency of the engine can also be increased by using less compressed air to cool the airfoils. The airfoil designer typically tries to maximize the cooling capability of the cooling air while also minimizing the amount of cooling air used in order to produce the highest level of efficiency increase.
FIG. 1 shows a prior art first stage turbine blade external pressure profile. As shown, the forward region of the pressure side surface experiences a high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than the pressure side. FIG. 2 shows a prior art turbine blade with a (1+5+1) forward flowing serpentine cooling circuit for the first stage blade. The flow path for the 5-pass serpentine flow circuit is shown in FIG. 3. For a forward flowing 5-pass serpentine cooling circuit used in the airfoil mid-chord region, the cooling air flows toward the leading edge and discharges into the high hot gas side pressure section of the pressure side represented by the dashed line in FIG. 1. In order to satisfy the backflow margin criteria, a high cooling supply pressure is needed for this particular design, and thus inducing a high leakage flow. In other words, the first leg of the 5-pass serpentine circuit has the highest pressure, and subsequent legs have a reduced pressure due to the travel of the cooling air through the passages. The last leg (fifth leg) in the serpentine circuit will therefore have the lowest pressure. The last leg in the FIG. 2 prior art circuit is also located at the highest external gas flow pressure. Thus, in order to prevent the hot gas flow from ingesting into the airfoil, the pressure in the last leg of the serpentine must be higher than the external gas flow pressure at the last leg. The inlet pressure for the 5-pass serpentine flow circuit must therefore be increased higher in order to prevent this from occurring.
In this particular cooling circuit, the blade tip section is cooled with double tip turns in conjunction with local film cooling. Cooling air bled off from the 5-pass serpentine flow circuit thus reduces the cooling performance for the serpentine flow circuit. Independent cooling flow circuit is used to provide cooling for the airfoil leading and trailing edge.
As TBC (thermal barrier coating) technology improves, more industrial turbine blades are applied with thick or low conductivity TBC. Cooling flow demand is reduced as a result of improved TBC protection. As a result, there is not sufficient cooling flow for the design with the prior art 1+5+1 forward flowing serpentine circuit of FIGS. 2 and 3. cooling flow for the blade leading and trailing edges has to be combined with the mid-chord flow circuit to form a single 5-pass flow circuit. However, for the forward 5-pass flow circuit with total blade cooling flow, BFM (back flow margin) may become a design issue.